Aircraft and missile forebody flow control device and method of controlling flow

ABSTRACT

A forebody flow control system and more particularly an aircraft or missile flow control system for enhanced maneuverability and stabilization at high angles of attack. The present invention further relates to a method of operating the flow control system. In one embodiment, the present invention includes a missile or aircraft comprising an afterbody and a forebody; at least one deployable flow effector on the missile or aircraft forebody; at least one sensors each having a signal associated therewith, the at least one sensor being used for determining or estimating flow separation or side forces on the missile forebody; and a closed loop control system; wherein the closed loop control system is used for activating and deactivating the at least one deployable flow effector based on at least in part the signal of the at least one sensor.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of then co-pending U.S. patentapplication Ser. No. 11/800,606, filed on May 7, 2007, and which issuedas U.S. Pat. No. 7,977,615 on Jul. 12, 2011, and which was acontinuation of then co-pending U.S. patent application No. 10/766,225,which was filed on Jan. 28, 2004, and which issued as U.S. Pat. No.7,226,015 on Jun. 5, 2007, which was a continuation of then co-pendingU.S. patent application Ser. No. 10/336,117, which was filed on Jan. 3,2003, and which issued as U.S. Pat. No. 6,685,143 on Feb. 3, 2004.

STATEMENT REGARDING FEDERALLY-SPONSORED RESEARCH

The U.S. Government has a paid-up license in this invention and theright in limited circumstances to require the patent owner to licenseothers on reasonable terms provided for by the terms of grant numbersF33615-98-C-3006 and F33615-99-C-3008 awarded by the Department ofDefense, Air Force Research Laboratories (AFRL) at Wright Patterson AirForce Base.

BACKGROUND OF THE INVENTION 1. FIELD OF THE INVENTION

The present invention relates to a forebody flow control system and moreparticularly to aircraft or missile flow control system for enhancedmaneuverability and stabilization at high angles of attack. The presentinvention further relates to a method of operating the flow controlsystem.

2. TECHNICAL BACKGROUND

In numerous aeronautical applications it is desirable to control theflow across a surface. As fluid flows over a flow surface, like air overan aircraft or a missile fore body, it forms a fluid boundary layer atthe surface. The fluid boundary layer is a thin layer of viscous flowexhibiting certain pressure variations that affect the operation of theaircraft or a missile.

One of these variations is the separation and vortex induced phantom yawcaused by asymmetric vortex shedding on an aircraft or a missile at highangles of attack, even at zero angle of sideslip of. Large forces anddynamic out-of-plane loading on the aircraft or missile occur at anglesof attack ranging from 30 to 60 degrees. It is known that theout-of-plane loading results from micro-asymmetries on the surface ofthe nose of the aircraft or missile such as dents, cracks in the paintand other microscopic imperfections near the tip of the nose. It hasalso been known that these asymmetries are affected by the bluntness ofthe forebody, Reynolds Number; roll angle, and the angle of attack. Athigh angles of attack, these side forces (yaw) are especially pronounceddue to ineffectiveness of the traditional flight control surfaces. Sideforces resulting from these asymmetries adversely affect the missile oraircraft's performance and significantly limit their flight envelope.

The demand for better control of missiles or aircraft at high angles ofattack has led to a number of approaches for control of these sideforces. Flow control devices have been employed to control andcounteract these side forces. These flow control devices are eitherpassive or active. Passive flow control devices have included geometricchanges to the forebody structure such as nose bluntness, strakes,boundary layer strips, vane vortex generators and rotating nose tips tocontrol the asymmetric vortices off the forebody. These passive flowcontrol techniques are effective to some extent in alleviating theseside forces, but at the same time limit the performance of the aircraftor missile by increasing the drag. Active flow control devices haveincluded jet blowing, unsteady bleed, suction, blowing and deployableflow effectors to control the asymmetric vortices off the fore body.These active flow control techniques are (as with passive devices) alsoeffective to some extent in alleviating these side forces, but also notoptimized (as with passive devices), because they operate in anopen-loop mode with no sensor feedback, at the same time limit theperformance of the aircraft or missile by increasing the drag.

In view of the foregoing disadvantages with presently available passiveor active flow control systems and methods for controlling flowasymmetries on a missile or an aircraft, it has become desirable todevelop a missile or aircraft forebody flow control system that controlsboth the magnitude and direction of these side forces (and further theaircraft or missile maneuverability), and can be deactivated when notrequired in order to reduce drag.

SUMMARY OF THE INVENTION

The present invention relates to a forebody flow control system and moreparticularly to aircraft or missile flow control system for enhancedmaneuverability and stabilization at high angles of attack. The presentinvention further relates to a method of operating the flow controlsystem.

In one embodiment, the present invention includes a missile or aircraftcomprising an afterbody and a forebody; at least one deployable floweffector on the missile or aircraft forebody; at least one sensor havinga signal associated therewith, the at least one sensor being positionedto detect flow separation on the missile or aircraft forebody; and aclosed loop control system; wherein the closed loop control system isused for activating and deactivating the at least one deployable floweffector based on at least in part the signal of the at least onesensor.

In another embodiment, the present invention includes a flow controlsystem for a missile or aircraft forebody comprising at least oneactivatable flow effectors; at least one sensor having a signal, the atleast one sensor being positioned to detect flow separation on themissile or aircraft forebody; an inertial measurement unit having anoutput; and a closed loop control system; wherein the closed loopcontrol system is used for activating and deactivating the at least oneflow effector based on at least in part the signal of the at least onesensor and the output of the inertial measurement unit.

In still another embodiment, the present invention includes a method ofstabilization for a missile or aircraft forebody comprising the steps ofestimating or determining side forces on a missile or an aircraftforebody based at least in part on a signal from at least one sensor,the at least one sensor being positioned to detect flow separation onthe missile or aircraft forebody; the missile or aircraft forebodyfurther comprising at least one flow effector and a closed loop controlsystem for controlling the flow effectors; activating the at least oneflow effectors to counteract the side forces by oscillation of the atleast one flow effector with the closed loop controller based on atleast in part the signal of the at least sensor; and re-estimating ordetermining side forces on the missile or aircraft forebody based atleast in part on a signal from the at least one sensor; and deactivatingthe at least one flow effector in response to reduced or changed sideforces.

Additional features and advantages of the invention will be set forth inthe detailed description which follows, and in part will be readilyapparent to those skilled in the art from that description or recognizedby practicing the invention as described herein, including the detaileddescription which follows, the claims, as well as the appended drawings.

It is to be understood that both the foregoing general description andthe following detailed description are merely exemplary of theinvention, and are intended to provide an overview or framework forunderstanding the nature and character of the invention as it isclaimed. The accompanying drawings are included to provide a furtherunderstanding of the invention, and are incorporated in and constitute apart of this specification. The drawings illustrate various embodimentsof the invention; and together with the description serve to explain theprinciples and operation of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1. Schematic view of one embodiment of a missile having a forebodywith flow effectors and sensors mounted therein.

FIG. 2. Schematic view of one embodiment of an aircraft forebody withflow effectors and sensors mounted therein.

FIG. 3. a) Perspective view of one embodiment of the forebody section ofa missile or aircraft having flow effectors and sensors mounted therein;b) Sectional view of forebody along plane A-A′ shown in FIG. 3a ).

FIG. 4. Perspective view of one embodiment of a module containing aco-located sensor, and a) a deployable flow effector (deployed) and b) adeployable flow effector (retracted).

FIG. 5. Sectional view of one embodiment of the forebody section of amissile or aircraft having flow effectors and sensors mounted therein.

FIG. 6. Sectional view of one embodiment of a deployable flow effector.

FIG. 7. Sectional view of deployable flow effector shapes.

FIG. 8. Sectional view of another embodiment of a deployable floweffector.

DESCRIPTION OF THE PREFERRED EMBODIMENT

The present invention relates to a forebody flow control system and moreparticularly to aircraft or missile flow control system for enhancedmaneuverability and stabilization preferably at high angles of attack.The forebody of the missile or aircraft for the present invention isdefined as the front half of the length of the missile or aircraft.Preferably, the forebody is the front 25% of the length of the missileor aircraft, and most preferably the forebody is the nose of the missileor aircraft. The nose of the missile or aircraft is the cone shapedleading edge. The activatable flow effectors of the present inventioninclude but are not limited to active vortex generators, which aredeployable or create pressure active regions by suction or air pressure.The present invention further relates to a method of operating the flowcontrol system.

The flow control system for stabilization and maneuverability of themissile or aircraft forebody relies on the effectiveness of theactivatable flow effectors in generating on-demand side forces aroundthe missile or aircraft forebody to create the desired stabilization ormaneuverability effect. The activatable flow effectors of the presentinvention are active micro-vortex generators that effectively controlthe pressure distribution along the forebody of the missile or aircraft,yielding large side forces and yawing moment for control of yaw oneither side of the forebody. The activatable flow effectors of thepresent invention preferably are deployable flow effectors or othertypes of micro-vortex generators. Activatable flow effectors of thepresent invention are flow effectors that are activated to generatefluid flow disturbances in the vicinity of the flow effector, and thatcan be deactivated when not needed. Preferably, the activatable floweffectors of the present invention can be activated and deactivated athigh frequencies. Further preferably, the activatable flow effectors arecapable of being cycled at frequencies of at least about 1 Hz, morepreferably at frequencies of at least about 20 Hz, even more preferablyat frequencies of at least about 60 Hz, and most preferably atfrequencies of at least about 100 Hz. Deployable flow effectors aredescribed in more detail in the various embodiments in the Figuresbelow. The frequencies at which the flow effectors of the presentinvention are cycled may be determined based in part on a number offactors including but not limited to autopilot frequency responsecharacteristics, missile or aircraft dynamics, and missile or aircraftenvironmental conditions. Some of the other types of flow effectors notshown in the Figures (but described in more detail in U.S. Pat. No.6,302,360 Bl to Ng which is herein incorporated by reference) includespaced apart valves that are positioned at inlets of a vacuum orpressure chamber, or are connected by pneumatics to a vacuum or pressuresource. Preferably, the valves contain a flap that operates to open andclose the valves as directed by electrostatic forces. Other valveconfigurations can also be used. When the valves are opened, thenegative pressure from the vacuum chamber or source causes withdrawal ofair from the surface of the missile or aircraft forebody through thesurface orifices. Therefore, it can be seen that the opening of thevalves causes the pressure active region to generate a net inflow of airfrom the upper flow of air traveling across the surface of the missileor aircraft forebody. This net inflow of air causes a disturbance in theupper flow, resulting in the generation of vortices, which actbeneficially to stabilize the airflow around the forebody surface of themissile or aircraft, or to create commanded side forces on the missileor aircraft forebody to improve maneuverability. Similarly, when thevalves are open to a positive pressure chamber or source, a net outflowof air is generated resulting in the generation of vortices, which alsoact beneficially to reattach the air flow to the forebody surface of themissile or aircraft. For purposes of this invention flow effectorsinclude any type of device or article known to those skilled in the artor discovered at a later point that is used to assist in thereattachment of airflow to a missile or aircraft's surface. Preferably,the flow effectors of the present invention are deployable floweffectors. Further preferably, the missile or aircraft of the presentinvention has at least about 4 activatable flow effectors, morepreferably at least about 6 activatable flow effectors and mostpreferably at least about 8 activatable flow effectors.

Referring now to FIG. 1, there is shown a schematic view of oneembodiment of a missile 10 having a forebody 18 and an afterbody 13. Theforebody 18 having at least one activatable flow effector 12. Theforebody further having at least one sensor 14. The sensor beingpositioned to detect flow separation from the flow surface 16 on themissile 10 forebody 18. The forebody 18 of this specific embodimenthaving a number of flow effectors 12 and sensors 14 mounted in theforebody 18 (or nose) therein. Furthermore in this specific embodiment,the individual flow effectors 12 and individual sensors 14 are in closeproximity with respect to each other. The fluid boundary layer is a thinlayer of viscous flow exhibiting certain pressure variationcharacteristics and fluid dynamics that affect the operation of the flowsurface 16. The fluid is generally air. The flow surface 16 for purposesof the present invention is the forebody of a missile or an aircraft.FIG. 2 is a schematic view of one embodiment of an aircraft 20 adaptedwith the vortex generating system 22 of the present invention. Theairplane can be any type of aircraft, including commercial, military andspace vehicles. The aircraft 22 includes a fuselage 21, a tail 23, wings24, forebody (nose) 18 and jet engines 26. In this specific embodiment,the individual flow effectors 12 and individual sensors 14 are alsomounted in close proximity with respect to each other on the forebody 18of the aircraft 20. Under certain conditions such as high angles ofattack, the missile 10 in FIG. 1 and the aircraft 20 in FIG. 2 mayexperience fluid boundary layer separation.

The sensor(s) of the present invention include but are not limited to adynamic pressure sensor, shear stress sensor (hot film anemometer, adirect measurement floating-element shear stress sensor), inertialmeasurement unit or system, and other sensors known to those skilled inthe art whose signal could be used to estimate or determine flowseparation on the surface of the missile or aircraft. The sensors of thepresent invention are used to determine or estimate flow separation. Aninertial measurement unit for example is a sensor, which to would notdirectly measure flow separation, but could be used to estimate orpredict separation. The preferred sensor of the present invention is apressure sensor. The pressure sensor is used to sense flow separation.The pressure sensor can be any type of sensor suitable for measuring thepressure at the flow surface. The pressure sensor can for example be apiezoelectric device, which generates an electric signal in response toa sensed pressure, a shape memory alloy device, or any other pressuresensor or transducer known to those skilled in the art. Preferably, theratio of flow effectors to sensors is less than about 3:1, morepreferably less than or equal to 2:1, and most preferably less than orequal to 1:1. The higher the concentration of pressure sensors to floweffectors the more redundancy can be built into the system utilizing thepresent invention. Most preferably the sensor is a flush, surfacemounted diaphragm type pressure sensor. The at least one sensor 14having a signal which is used at least in part by a controller (notshown) to activate and deactivate the at least one flow effector 12.

In addition to flow separation sensors, various embodiments of thepresent invention may also include a means for determining the relativespatial orientation of the flow effectors and/or sensors with respect tothe flow separation on the missile or aircraft body. This means wouldinclude utilizing the output of an inertial measurement unit and othersystems, which could be used to determine the missile or aircraftorientation with respect to this flow separation. An inertialmeasurement unit provides six-degree-of-freedom motion sensing forapplications such as navigation and control systems. Angular rate andacceleration are measured about three orthogonal axes.

FIG. 3a ) is a perspective view of one embodiment of the forebodysection of a missile or aircraft having activatable flow effectors 12and sensors 14 mounted therein. The missile or aircraft forebody of thepresent invention can be designed with asymmetries in the forebody (15)to provide better stability or control with the present clow controlsystem. Boundary layer separation at the missile or aircraft forebody iscaused by a combination of the viscous forces within the fluid boundarylayer and an adverse, pressure gradient over the flow surface 16.Controlling fluid boundary layer dynamics not only provides an overallbenefit to the operation of the flow surface but also counteracts andcontrols fluid boundary layer separation. Due to the geometricallyslender body or micro-asymmetries at the nose of a missile or anaircraft, boundary layer flow separation of the fluid flow 28 at theflow surface 16 at high angles of attack 27 may result in large sideforces and dynamic out-of-plane loading resulting in a yawing moment ofa missile or an aircraft. In other words, at high angles ofattack)(>15°) of the forebody of an aircraft or missile there may besome degree of asymmetric vortex shedding. Asymmetric vortex shedding iscaused by fluid passing over the missile or aircraft and separating onone-side of the missile or aircraft prior to separation on the other (orto a greater extent). One of the objects of the present invention is tostabilize, control and/or create side forces to improve the stabilityand maneuverability of a missile or an aircraft. High angles of attack27 are represented by theta (φ) 27. High angles of attack are preferablya theta (φ) 27 of at least about 20°, and more preferably a theta (φ)from about 30° to about 60°. FIG. 3b ) is a sectional view of sectionA-A′ of a missile or aircraft forebody 18 as shown in FIG. 3a ). FIG. 3b) shows the fluid flow 28 around a missile or aircraft forebody 18 at asection A-A′ in the proximity of the activatable flow effectors 12, andthe resultant flow separation prior to activation of the flow effectors12.

In FIG. 4, there is shown a perspective view of one embodiment of amodule containing a co-located sensor, and a) an activatable, deployableflow effector (deployed) and b) an activatable, deployable flow effector(retracted). In this particular embodiment, the module 32 contains anactivatable, deployable flow effector 12 and a pressure sensor 14. Theactivatable, deployable flow effector 12 being capable of being deployedinto and retracted from, respectively, the fluid boundary layer flowingover the flow surface of the missile or aircraft forebody wherein themodule 32 is employed. The deploying and retracting can be accomplishedusing any device such as pneumatic pressure, hydraulic pressure, vacuum,a mechanical device such as a solenoid valve, a microelectromechanicaldevice, any combination thereof or the like. The module 32 may or maynot include a controller (not shown) internal to the module. Thepressure sensor 14 is connected to the controller (not shown). If thecontroller (not shown) is not internal to the module 32 then the module32 preferably further comprises a link between pressure sensor 14 andthe controller, and another link between the controller (not shown) anddeploying means (not shown). The controller (not shown) is programmed tooperate the deploying and retracting means in response to specificpressure conditions sensed at the flow surface 16. The controller (notshown) can be any device such as a computer, suitable for gatheringinformation from the pressure sensors 14, and directing the activationof the activatable flow effectors 12. Where a number of activatable floweffectors 12 and/or pressure sensors 14 (or modules 32) are employed,the controller (or controllers) (not shown) can be programmed andconnected to integrate each of the activatable flow effectors 12,pressure sensors 14 and modules 32 so that the output from all of theregions will be coordinated to enhance and possibly optimize thestabilization and maneuverability of a missile or an aircraft forebody.Specific patterns of deployment and/or retraction of the flow effectors12 can be determined to handle a variety of routine events and alsoincorporated into the control scheme.

FIG. 5 is a sectional view of one embodiment of the forebody (nose)section of a missile or aircraft having a flow effector 12 and sensor 14mounted therein. In FIG. 5, the two activatable flow effectors 12 shownin this cross-section are movably attached by an attachment means, i.e.,a hinge 91, to a base structure 82. The activatable flow effectors 12are deployable flow effectors. The activatable flow effectors 12 arefurther movably attached to a piston 84. The piston 84 moves within acylinder 86 in response to a pressure source (not shown) applied via apneumatic system (not shown) against an elastomeric sheet 81 to move thepistons 84 and in return to deploy and retract the flow effectors 12.The piston 84 also is connected to a biasing means 87, i.e., a spring,to return the piston 84 to its original position upon removing thepressure source, and therefore retracting the deployable flow effector12. In this particular embodiment, the pressure is applied to the piston84 via a pressure inlet/outlet 88. Also shown in this particularembodiment are seals in the form of O-rings 83 to seal the pneumaticsystem (not shown) of the pressure source (not shown); and two sensors14. The sensors 14 are connected via leads to a controller (not shown).The pressure source (not shown) is also connected to the controller (notshown).

FIG. 6 is a sectional, detailed view of a module 32 (as shown in FIG. 4)with an activatable deployed flow effector 12. In FIG. 6, the floweffector 12 is movably attached to the upper portion 48 of the housing46 of the module 32 and is attached to the lower portion 50 of thehousing 46 of the module 32 by at least two fasteners 40. The upperportion 48 of the housing 46 mates with the lower portion 50 with asealing ring (not shown) and a sealable, flexible element 44 therebetween. The flow effector 12 is deployed by pressure being applied tothe flexible element 44. The flow effector 12 has a biasing means (aspring) 41 which attaches at one end to the upper portion 48 of thehousing 46 and at the other end to the base 54 of flow effector 12.Directly beneath the flow effector 12 is a valve 43, which opens andcloses to allow for the application of fluid or gas pressure from apressure source not shown to be applied to the flexible element 44through a pneumatic pathway 52. A pressure sensor 14 senses fluid flowat or near the surface over which the fluid is flowing. Preferably thepressure sensor is at the surface of the airfoil, and most preferably itis flush with such surface. The pressure sensor 14 can be any pressuresensor but advantageously is a microelectromechanical (MEMS) based orpiezoelectric based sensor. The sensor transmits a signal, in this casea voltage but it is understood to one skilled in the art that the signalcan be other than voltage, including, but not limited to, current,pressure, hydraulic or optical. The signal corresponds to the pressureit senses.

The pressure sensor 14 is connected to a controller 42 internal to themodule 12 (or optionally external to the module). The controller 42 canbe for example a proportional-integral-derivative (PID) controller, anadaptive predictive controller, or an adaptive predictive feedbackcontroller. The controller of the present invention is preferably aclosed loop control system. The controller can be used to minimize sideforces or to create commanded side forces on the missile or aircraftforebody. The pressure sensor transmits a signal to the controller 42through the electrical connection 38 (in practical application, multiplepressure sensors 14 send multiple signals to the controller 42). Thecontroller 42 processes the signals to determine, through mathematicalmodeling, the dynamics of the flow surface. Such dynamics includeboundary layer separation and stall. It is the predictive ability of thecontroller 42, which provides for this function and expands this systemfrom being merely responsive. This is especially advantageous fordynamic systems, which are nonlinear and time varying and operating inchallenging environments. The controller 42 produces an output signal toa monitor, recorder, alarm and/or any peripheral device for alarming,monitoring, or in some manner, affecting or precluding the dynamics uponits incipience. Advantageously, the controller 42 is the ORICA™controller, an extended horizon, adaptive, predictive controller,produced by Orbital Research, Inc. and patented under U.S. Pat. No.5,424,942, which is incorporated herein by reference. Under certainconditions, the controller 42 (or optionally an external controller)which is connected via electrical connection 46 to the valve 43 causesthe valve 43 to open thereby resulting in the deployment of the floweffector(s) 12. Preferably, the pressure source (or other deploymentand/or retraction means) is internal to the module 12. The sealable,flexible element 44 referred to above can be made of a single polymer ora combination of polymers. The pressure source can be air bled from anaircraft turbine engine, a pressurized gas cartridge, or pressurizedfluid. The biasing means is employed to urge the sealable, flexibleelement 44 towards its quiescent state after pressure is removed orreduced. The biasing means can be any device or spring like means, suchas vacuum or pressure, mechanical or electromechanical device.

The deployable portion of the activatable, deployable flow effectorsshown in the previous Figures are small mechanical tabs preferably madefrom epoxy glass-fabric, and deactivate to assume a position underneaththe skin surface of the missile or aircraft in their retracted state.Several examples of various embodiments of the flow effectors are shownin FIG. 7. a, b, c and d. These cross-sectional views demonstrate thatrectangular 72, triangular 74, irregular 76, semi-circular 78, andsquare not shown can be used. The present invention is, however, notlimited to these shapes and it is envisioned that any shape of floweffector known presently or conceived of in the future by those skilledin the art may be used. Upon controlled activation, the flow effectors(deployable or other) manipulate the forebody of the missile oraircraft's vortical flow field to generate the desired side forces oryawing moment. Single flow effectors or combinations of flow effectorscan be activated either statically or cycled at a varying frequency(oscillated) to obtain a desired side force or yawing moment. Varyingfrequency or oscillation of the flow effectors includes but is notlimited to pulse width modulation or other techniques known to thoseskilled in the art.

FIG. 8 is a sectional view of another embodiment of a deployable floweffector. In FIG. 7, the activatable flow effectors 12 are deployableflow effectors. The flow effectors 12 are further movably attached to acamshaft 94. The camshaft 94 moves in response to an electric motor 96to deploy and retract the flow effector 12. The motor is connected to acontroller 42. The controller 42 activates and deactivates thedeployable flow effector in response to at least in part the signal fromthe sensor 14.

It will be apparent to those skilled in the art that variousmodifications and variations can be made to the present inventionwithout departing from the spirit and scope of the invention. Thus, itis intended that the present invention cover the modifications andvariations of this invention provided they cone within the scope of theappended claims and their equivalents.

What is claimed:
 1. A method of stabilization and maneuvering a missileor aircraft comprising the steps of a. estimating or determining sideforces on a missile or an aircraft forebody based at least in part on asignal from at least one sensor; the missile or aircraft furthercomprising at least one deployable and retractable flow effector on themissile or aircraft forebody and a closed loop control system forcontrolling the at least one deployable and retractable flow effector atleast in part by indirectly estimating or determining flow separation,or estimating or determining side forces on the forebody based on thesignal from at least one sensor; b. activating the at least onedeployable and retractable flow effector to counteract or change theside forces by activation of the at least one deployable and retractableflow effector with the closed loop controller based on at least in partthe signal of the at least one sensor; c. re-estimating or determiningside forces on the missile or aircraft forebody based at least in parton a signal from the at least one sensor; and d. deactivating the atleast one deployable and retractable flow effector by retracting theflow effector to a position at or beneath the surface of the missile oraircraft in response to reduced or changed side forces.
 2. The method ofstabilization or maneuvering in claim 1, wherein the at least one floweffector is activated by deploying the at least one flow effector. 3.The method of stabilization or maneuvering in claim 1, wherein themissile or aircraft forebody comprises at least six flow effectorswherein the at least six flow effectors are positioned and separatedsubstantially equidistantly about a center of the forebody of themissile or aircraft, and the forebody is the front 25% of the length ofthe missile or aircraft.
 4. The method of stabilization or maneuveringin claim 1, wherein the forebody of the missile or aircraft is designedwith asymmetries in the forebody.
 5. The method of stabilization ormaneuvering in claim 1, wherein the at least one flow effector isactivated at angles of attack of the missile or aircraft forebody ofgreater than about 20 degrees to counteract out-of-plane loading on themissile or aircraft caused by asymmetries on the forebody.
 6. The methodof stabilization or maneuvering in claim 1, wherein the aircraft ormissile further comprises an inertial measurement unit with an output,and the output of the inertial measurement unit is used to determine themissile or aircraft orientation with respect to a flow separationdetermined by the at least one sensor.
 7. A method of stabilization ormaneuvering a missile or aircraft comprising the steps of a. estimatingor determining side forces on a missile or an aircraft forebody based atleast in part on a signal from at least one sensor; the missile oraircraft further comprising at least one flow effector on the missile oraircraft forebody, the forebody being the front 25% of the length of themissile or aircraft, and a closed loop control system for controllingthe at least one flow effector at least in part by indirectly estimatingor determining flow separation, or estimating or determining side forceson the missile forebody based the signal from at least one sensor; b.activating the at least one flow effector to counteract or change theside forces by activation of the at least one flow effector with theclosed loop controller based on at least in part the signal of the atleast one sensor; and c. re-estimating or determining side forces on themissile or aircraft forebody based at least in part on a signal from theat least one sensor.
 8. The method of stabilization or maneuvering inclaim 7, wherein the at least one flow effector is positioned on thenose of the missile or aircraft.
 9. The method of stabilization ormaneuvering in claim 8, wherein the at least one flow effector isdeployable and retractable.
 10. The method of stabilization ormaneuvering in claim 9, wherein the at least one flow effector isactivated by deploying the at least one flow effector.
 11. The method ofstabilization or maneuvering in claim 7, wherein the missile or aircraftforebody comprises at least six flow effectors wherein the at least sixflow effectors are positioned and separated substantially equidistantlyabout a center of the forebody the missile or aircraft.
 12. The methodof stabilization or maneuvering in claim 7, wherein the forebody of themissile is designed with asymmetries in the forebody.
 13. The method ofstabilization or maneuvering in claim 7, wherein the least one floweffector is activated at angles of attack of the missile or aircraftforebody of greater than about 20 degrees to counteract out-of-planeloading on the missile or aircraft caused by asymmetries on theforebody.
 14. The method of stabilization or maneuvering in claim 7,wherein the aircraft or missile further comprises an inertialmeasurement unit with an output, and the output of the inertialmeasurement unit is used to determine the missile or aircraftorientation with respect to a flow separation determined by the at leastone sensor.
 15. A method of stabilization or maneuvering a missile oraircraft comprising the steps of a. estimating or determining sideforces on a missile or an aircraft forebody based at least in part on asignal from at least one sensor; the missile or aircraft furthercomprising at least one flow effector on the missile or aircraftforebody and a closed loop control system for controlling the at leastone flow effector at least in part by indirectly estimating ordetermining flow separation, or estimating or determining side forces onthe missile forebody based on the signal from at least one sensor; b.activating the at least one flow effector to counteract or change theside forces by activation of the at least one flow effector with theclosed loop controller based on at least in part the signal of the atleast one sensor; and re-estimating or determining side forces on themissile or aircraft forebody based at least in part on a signal from theat least one sensor; wherein the at least one flow effector is activatedat angles of attack of the missile or aircraft forebody of greater thanabout 20 degrees to counteract out-of-plane loading on the missile oraircraft caused by asymmetries on the forebody of the missile oraircraft.
 16. The method of stabilization or maneuvering in claim 15,wherein the at least one flow effector is deployable and retractable.17. The method of stabilization or maneuvering in claim 15, wherein theat least one flow effector is activated by deploying the at least oneflow effector.
 18. The method of stabilization or maneuvering in claim15, wherein the missile or aircraft forebody comprises at least six floweffectors wherein the at least six now effectors are positioned andseparated substantially equidistantly about a center of the forebody ofthe missile or aircraft, and the forebody is the front 25% of the lengthof the missile or aircraft.
 19. The method of stabilization ormaneuvering in claim 15, wherein the forebody of the missile or aircraftis designed asymmetries in the forebody.
 20. The method of stabilizationor maneuvering in claim 15, wherein the aircraft or missile furthercomprises an inertial measurement unit with an output, and the output ofthe inertial measurement unit used to determine the missile or aircraftorientation with respect to a flow separation determined by the at leastone sensor.